1. Field of the Invention
The present invention is directed to a method and composition for protecting spacecraft from electrostatic discharge and thermal effects, and, more particularly, to a polymer composition incorporating lithium salts and optionally tantalum pentoxide and the method of use of such a composition.
2. Description of Related Art
Satellites and spacecraft must be able to withstand the stringent conditions imposed by the space environment, such as temperature extremes (e.g. -156.degree. C. or -250.degree. F. to 121.degree. C. or 250.degree. F.) for extended periods of time and space vacuum. In particular, an on going problem for spacecraft has been the difficulty of providing thermal control in order to maintain the electronics and batteries on the spacecraft at or near room temperature to optimize their performance. In addition, some structural components used in spacecraft are formed of organic materials which can be damaged when exposed to a temperature above 200.degree. F. (93.degree. C.), for example. Consequently, these components must be cooled in order to avoid structural damage.
The central problem in thermal control for spacecraft is the absence of convection in the space environment, where objects cool only by radiation, which is far less effective in transferring heat than conduction. This problem is further complicated by the fact that the heat load due to solar exposure is very high so that temperature extremes are frequently encountered.
Any object constantly emits electromagnetic radiation in all directions. If the object is below about 1000.degree. F. (538.degree. C.), the energy is infrared, which is invisible to the human eye. The amount of energy emitted is directly related to a parameter called "emissivity," normally denoted as .epsilon.. When a surface is exposed to electromagnetic radiation, such as sunlight, the radiation is partially absorbed, partially reflected, and if the body is transparent or translucent, partially transmitted. The relation between the three energy components is given by the conservation equation EQU .alpha.+.rho.+.tau.=1
where
.alpha.=absorptivity: the fraction of the incident radiation absorbed by the body PA1 .rho.=reflectivity: the fraction of the incident radiation reflected from the surface PA1 .tau.=transmissivity: the fraction of the incident radiation which passes through the body, i.e., is transmitted PA1 Y=H, C1 to C4 alkyl, OH, or halogen, PA1 Z=H or halogen, PA1 R.sub.1 and R.sub.2 are each a substituent compatible with the polymer;
The relative magnitudes of .alpha., .rho., and .tau. depend not only on the material properties, temperature, and geometry, but also on the wavelength or spectrum of the radiation. For most of the materials of present interest, .tau. is small (less than 0.01), and can be considered zero. In general, a material with low .alpha. is relatively unaffected by sunlight; a surface with high .epsilon. is a good heat rejector; and a surface with a low .alpha./.epsilon. ratio tends to stay cold when illuminated.
Thermal control of spacecraft has been achieved by using films or multilayer blankets of a dielectric material, such as Kapton (a polyimide which is a registered trademark of E. I. DuPont) or Teflon (a polytetrafluoroethylene which is a registered trademark of E. I. DuPont). These films are coated with aluminum or silver on the inner surface of the film to achieve high reflectance and low absorptivity. A thermal control blanket may be formed, for example, from several (e.g. 3 to 6) layers of Kapton film which have been aluminized on one surface and which are assembled such that contact between adjacent layers is minimized. The thermal control blanket is applied to the exterior surface of the part to be protected. However, the dielectric inner layers of such a thermal control blanket become charged when high energy particles in space penetrate the outer metallized layers and become fixed in the inner layers of the blanket. This charge can accumulate to a high level, at which time electrostatic discharge can occur, as discussed below.
A spacecraft in the space environment is exposed to numerous charged particles and radiation. Static charge builds up in portions of the spacecraft in the form of excess electrons. A voltage differential builds up between separate portions of the spacecraft and can reach levels sufficient to cause electrostatic discharge between various surfaces in the spacecraft or arcing to structural ground. These discharges can damage or degrade electronic circuits, which can produce such significant problems as a power outage, temporary loss of communication to the ground, or loss of system control functions. In addition, these static discharges can degrade the optical properties of thermal control surfaces.
Fortunately, the problem of static discharge is somewhat alleviated in some cases since radiation from the sun causes the emission of stored electrons, to thus neutralize the static charge buildup. Consequently, for a satellite which spins continuously, that is, a spin-stabilized satellite, static charge buildup is somewhat alleviated since only a small fraction of the satellite is in shadow long enough to build up electrostatic discharge and this charge is dissipated by solar photons in each spin cycle. However, for a satellite which does not spin continuously, that is, a body-stabilized satellite, one half of the satellite is always in shadow. The shadowed areas cannot photoemit stored electrons efficiently compared to the adjacent sun-illuminated areas. A voltage differential is thus produced and gives rise to electrostatic discharge events. In addition, the rear surfaces of the flat solar panels used in satellites cannot bleed off electrostatic charge at a rate greater than the geomagnetic plasma can charge them in a body-stabilized satellite.
At the present, attempts to reduce electrostatic discharge in satellites have included the use of outer coatings possessing surface conductivities only (indium tin oxide or germanium), or the use of filter pins in wire harness connectors. Filter pins in wire harness connectors have the disadvantage that they merely attenuate electrostatic discharge events, but do not prevent them. With respect to the thin coatings (100 to 2000 .ANG.), they are fragile, costly, and provide surface conductivities only. They do not provide protection from geomagnetic plasma storms since they can be easily penetrated by electrons with energies as low as 5 keV, allowing the underlying structure to electrostatically charge.
Thus, a need exists in the field of spacecraft and satellites for a material composition which can provide thermal control and at the same time protect against electrostatic discharge in space.